High performance gas turbine engines include, as depicted schematically in FIG. 1 along longitudinal axis 101, a compressor 4, a combustor 6, a high pressure turbine nozzle 10, and a high pressure turbine 50. Core air is compressed by the compressor and discharged in axial flow as high pressure air to the combustor where fuel is injected and ignited. The hot, pressurized gases, which in modern engines can be in the range of 2,000.degree. F., are allowed to expand through the high pressure turbine nozzle which directs the flow to turn the turbine, which is coupled by a rotor shaft 51 to drive the compressor. The core gases then exit the high pressure turbine providing energy downstream in the form of additional rotational energy extracted by additional and lower pressure turbine stages and/or thrust through an exhaust nozzle.
A portion of high pressure air can be bled from the compressor and used as high pressure cooling air to cool downstream apparatus, such as the combustor, nozzle, and turbine, and high pressure cavity purge air, but such high pressure cooling and cavity purge air consumes work from the turbine and is quite costly in terms of engine performance. Reducing the cooling and cavity purge air requirement allows a higher core air flow, reduces the energy expended by the turbine, increasing the energy available in the gas flow path.
In addition to using high pressure cooling air to protect the combustor and subsequent components from the effects of high temperature gases, such as high gas velocity oxidation and thermal fatigue, gas turbine engines can also employ protective coatings such as Thermal Barrier Coatings (TBC) to protect engine parts. Typically a TBC when applied to a metal substrate, protects the substrate from the effects of exposure to high temperature gases and act as an insulating layer between the hot gas flow path and the substrate. A thermal barrier coating system typically has multiple layers. Typically, such a system has at least a bond coat layer such as MCrAlY and a top coat such as a ceramic, like Yttria-stabilized zirconia layer, and may include additional layers.
A high pressure turbine nozzle, such as described in Aircraft Gas Turbine Engine Technology, 2d edition, (McGraw-Hill, 1979), pages 480-481 and incorporated herein by reference, typically comprises circumferentially adjacent paired vane segments, each vane extending radially outward from an inner to an outer band. The nozzle defines an annular core gas flowpath, turning the core flow to an angle for optimum performance of the turbine. The nozzle is segmented into the paired vanes to limit problems from differing thermal responses between the nozzle and supporting structure, creating gaps to allow for thermal growth of the segments. The nozzle uses high pressure cooling air for convection, impingement, and film cooling. Spline seals are placed in slots formed in the segments inner and outer bands circumferentially abutting surfaces to prevent leakage of hot core gases and uncontrolled loss of high pressure coolant air flow through these gaps.
The nozzle is positioned by seating the inner and outer bands of each nozzle segment in inner and outer support members, which transfer the pressure loads experienced by the nozzle to the engine casings and frame through the cold structures. The outer support member is not part of this invention. The inner band is typically mounted and retained to the inner nozzle support by means of bolts, pins or a combination of both at the inner band aft flange.
A gap exists at the interface between the high pressure nozzle inner band and turbine rotor blades. Any hot core gas leaking across this interface exits the working gas stream with a resultant loss of energy available from the gas turbine engine. Additionally, the turbine and nozzle structural members need to be protected from these hot core gases. To limit hot gas ingestion into this gap and to shield the inner nozzle support from high temperature gases, circumferential discourager seals are located in series, radially, typically including a nozzle inner band overhang, a high pressure turbine blade angel wing, and discourager seal(s) bolted to the nozzle support. Each discourager seal typically consists of circumferentially segmented members, creating split lines that permit leakage. A series of discourager seals create an air flow circuit with cavities in flow communication with each other. High pressure air that originated as high pressure cooling air can be vented to the rotor/stator interface, and used to purge these cavities of hot gases, the turbine acting to pump the air radially outward, adding sufficient energy to force the air through the discourager seals and to prevent ingestion of hot core gases. Such cavity purge air is a loss chargeable to the performance of the gas turbine engine and therefore the requirement for this flow should be minimized.
As the temperature of the cavity otherwise exposed to the inner nozzle support often exceeds the material capability of the support, discourager seals bolted to the support are typically thin, sheet metal seals that also act to insulate the support. Due to high thermal gradients, the seals are segmented to cut tangential stress, and are not integral with the support, which is typically not segmented, due to the differing thermal environment and thermal response characteristics. Additionally, many engines include a separate stationary seal between the rotor and the nozzle support which is also typically bolted in position.
Because the turbine rotates at high speeds while the nozzle assembly remains stationary, any structural protrusion into the rotor/stator cavity causes "windage losses." Such losses include the direct mechanical effect of work required to accelerate air to the speed of the rotor, which induces a drag effect on the rotor, and also the indirect effect of the resultant temperature rise in the purge air. High temperature air in the cavity needs to be avoided, thus requiring a higher purge flow. Every discontinuity in the flowpath environment causes windage losses, as the air is slowed and then must be reaccelerated. Segmentation of the discourager seal has been found to cause such losses, as does each bolthead or other threaded fastener exposed to the rotor/stator cavity.
One prior solution was to replace bolts with recessed Torx head screws and to protect the bolt heads with windage covers. However, because of the hostile temperature environment in this area, bolts and screws frequently become seized making disassembly without damage difficult. The parts count and complexity of assembly also made this solution unattractive.
Bolts or other fasteners which protrude into the rotor/stator cavity also affects how long the nozzle inner band overhang needs to be. The overhang is a hard area to cool and shortening the overhang would reduce the requirement for high pressure cooling air.
In addition to the above performance related problems encountered in using a segmented discourager seal and threaded fasteners to assemble the high pressure turbine nozzle, maintainability requirements for hot section structures such as the high pressure turbine nozzle that require periodic inspection and replacement necessitate an improved mounting arrangement that reduces the complexity and number of parts involved in assembly and disassembly.